WO2004031656A1 - Turbine a gaz - Google Patents
Turbine a gaz Download PDFInfo
- Publication number
- WO2004031656A1 WO2004031656A1 PCT/EP2003/009703 EP0309703W WO2004031656A1 WO 2004031656 A1 WO2004031656 A1 WO 2004031656A1 EP 0309703 W EP0309703 W EP 0309703W WO 2004031656 A1 WO2004031656 A1 WO 2004031656A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- combustion chamber
- coolant
- gas turbine
- tubes
- turbine
- Prior art date
Links
- 238000002485 combustion reaction Methods 0.000 claims abstract description 114
- 239000002826 coolant Substances 0.000 claims abstract description 96
- 230000007704 transition Effects 0.000 claims description 10
- 239000000463 material Substances 0.000 claims description 9
- 238000010276 construction Methods 0.000 abstract description 10
- 239000000446 fuel Substances 0.000 abstract description 6
- 238000009434 installation Methods 0.000 abstract description 2
- 239000012530 fluid Substances 0.000 abstract 1
- 238000001816 cooling Methods 0.000 description 31
- 238000013461 design Methods 0.000 description 11
- 238000004519 manufacturing process Methods 0.000 description 5
- 238000012423 maintenance Methods 0.000 description 3
- 230000008646 thermal stress Effects 0.000 description 3
- 238000010438 heat treatment Methods 0.000 description 2
- 238000005304 joining Methods 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000012544 monitoring process Methods 0.000 description 2
- 239000000523 sample Substances 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 238000007493 shaping process Methods 0.000 description 2
- 238000012546 transfer Methods 0.000 description 2
- 229910001018 Cast iron Inorganic materials 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 239000000919 ceramic Substances 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 238000003745 diagnosis Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000002349 favourable effect Effects 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 239000007788 liquid Substances 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 238000005192 partition Methods 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- 239000011241 protective layer Substances 0.000 description 1
- 230000011218 segmentation Effects 0.000 description 1
- 238000011144 upstream manufacturing Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/005—Combined with pressure or heat exchangers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00012—Details of sealing devices
Definitions
- the invention relates to a gas turbine having a combustion chamber in which a fuel supplied is reacted with combustion air supplied to produce a working medium.
- Gas turbines are used in many areas to drive generators or work machines.
- the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
- the fuel is burned in a number of burners, compressed air being supplied by an air compressor.
- the combustion of the fuel creates a working medium under high pressure at a high temperature.
- This working medium is fed into a turbine unit downstream of the respective burner, where it relaxes while working.
- Each burner can be assigned a separate combustion chamber, the working medium flowing out of the combustion chambers being able to be brought together in front of or in the turbine unit.
- the gas turbine can also be designed in a so-called annular combustion chamber design, in which a plurality, in particular all, of the burners open into a common, usually annular, combustion chamber.
- a cooling system constructed in this way has the disadvantage that the construction of the combustion chamber and the cooling system is very complex.
- the actual combustion chamber wall is a separate one on the outside
- Cooling system which in turn must be attached from the outside, assigned.
- the manufacturing process of such a combustion chamber can thus be very costly and labor-intensive, since many individual parts and joining processes are necessary for the production. This also leads to an increased susceptibility to errors during manufacture and operation in the gas turbine. Maintenance and repair work is also made more difficult by the complicated combustion chamber wall construction.
- the invention is therefore based on the object of specifying a gas turbine which has a particularly high efficiency even with a simple construction.
- the wall of the combustion chamber is formed from coolant tubes.
- the invention is based on the consideration that the gas turbine should be designed to ensure particularly high efficiency for particularly high media temperatures. In order to keep the susceptibility to errors low, particularly reliable cooling of the thermally loaded components, in particular the combustion chamber, should be ensured. This can be achieved with comparatively little effort, in that the combustion chamber wall, on the one hand, is also designed to be coolable and, on the other hand, is constructed from molded parts which are kept relatively simple and flexible.
- the peripheral wall of the combustion chamber or combustion chamber wall is suitably constructed from pipes.
- cooling air is provided as the coolant, which, after passing through the coolant tubes, can be supplied to the combustion chamber as additional combustion air preheated as a result of the combustion chamber cooling.
- the combustion chamber wall, the coolant tubes are advantageously made of cast material, in other words, each form a cast part.
- Another advantage of this choice of material is that reliable thermal insulation is made possible in a particularly simple manner by providing a suitable coating of the cast material with a ceramic protective layer.
- coolant tubes In order to keep the coolant tubes particularly insensitive to thermal stresses and thus particularly robust, they are designed in an advantageous embodiment with a trapezoidal cross section.
- This cross-sectional shape has a particularly high thermal elasticity which, even when the individual circumferential segments of the respective tube are heated to very different degrees, only leads to low thermal stresses between cold and warmer regions of the tube, so that a long service life of the coolant pipes can be achieved.
- the coolant tubes are expediently attached to carrier rings oriented in the circumferential direction of the combustion chamber. Due to their position and shape, these carrier rings determine the shape of the annular space of the combustion chamber formed by the coolant tubes. In the manner of a self-supporting structure, the production of a mechanically stable combustion chamber structure is made possible by using only a small number of further components in addition to the actual pipes.
- the coolant tubes are expediently attached to the carrier rings by means of cooled screws.
- the fastening of the coolant pipes by means of screws allows a particularly time-saving assembly or disassembly of one or more coolant pipes from the hot gas side, that is, without having to disassemble the combustion chamber.
- the carrier rings are advantageously connected to one another by a number of longitudinal ribs in addition to the actual coolant tubes.
- the carrier rings and longitudinal ribs are preferably welded together, so that the rings and ribs form a welded supporting body.
- a particularly high degree of flexibility in shaping the combustion chamber which in particular allows flow conditions in the working medium to be taken into account in the combustion chamber, while at the same time having a sufficient length and
- the shape of the coolant tubes can be guaranteed can be achieved by the cooling tubes expediently consisting of two or more tube segments connected to one another in their longitudinal direction.
- the advantage of segmenting the pipes can be, in particular, that production-technical difficulties in producing coolant pipes from cast iron with a sufficient length and appropriate shape are avoided.
- each segment preferably has an associated transition or connecting piece at its respective tube end.
- the transition pieces are expediently designed for easy connection to one another.
- the transition pieces are in particular selected such that segments can be connected by means of a plug connection. If there is a trapezoidal cross section of the coolant pipes, the cross section of the transition piece is expediently selected such that it changes to a circular cross section right up to the connection point or to the respective pipe segment end. Such a circular end cross-section particularly enables simple machining options for a precisely fitting connection with the adjoining pipe segment.
- this cooling system has the advantage that it is integrated into the wall structure of the combustion chamber and therefore only requires a few additional parts for the construction of the cooling system.
- this cooling system due to the comparatively straight discharge of the coolant, there is only a slight loss of coolant pressure. This has the advantage that a high efficiency of the turbine is also favored on the coolant side.
- the heat input into the coolant is advantageously recovered for the actual energy conversion process in the gas turbine.
- the cooling air heated during the combustion chamber cooling and used as a coolant is advantageously fed into the combustion chamber, the preheated cooling air being able to serve as exclusive or additional combustion air.
- each coolant tube is preferably connected on the output side to a collecting space, which in turn is connected upstream of the combustion chamber on the air side.
- the coolant can be mixed with the rest of the compressor mass flow via a throttle device and fed to the combustion process.
- a more even flow can be achieved by advantageously assigning such a collecting space to each burner, with the same amount of cooling air or coolant flowing to each collecting space by design.
- each burner is preferably connected to a collection space, each collection space being connected to the same number of coolant tubes.
- This arrangement has the particular advantage that approximately the same amount of recirculated cooling air is fed to each burner.
- the combustion chamber is designed as an annular combustion chamber, the combustion chamber therefore has a particularly uniform combustion process.
- the advantages achieved by the invention are, in particular, that the design of the combustion chamber wall as a plurality of interconnected coolant tubes provided for the flow of a coolant, in particular cooling air, enables particularly reliable combustion chamber cooling with a simple structure.
- the integration of the coolant pipes in a self-supporting combustion chamber structure, in particular by means of the carrier rings, also enables comparatively simple interchangeability of individual pipes requiring maintenance, but due to the flexibility that can be achieved via the pipe construction, replacement of existing combustion chamber structures in already existing gas turbines is also possible in a simple manner is.
- the construction of the combustion chamber from pipes is comparatively stable and insensitive to vibrations of the combustion chamber wall, since the coolant pipes stiffen and solidify the annulus.
- the basic flexibility in terms of shape and component selection achieved through the construction of the combustion chamber wall from tubular elements also allows, in particular, the attachment of probes or monitoring sensors for monitoring and / or diagnosis of the actual combustion process in the combustion chamber, in particular through the targeted use of specifically modified pipes allow, for example, the passage of suitable probes from the outside into the interior of the combustion chamber.
- FIG. 1 shows a half section through a gas turbine
- FIG. 2 shows in longitudinal section a segment of the combustion chamber of the gas turbine according to FIG. 1, and
- FIGS. 3a to c each show a cross section of a section of the combustion chamber wall according to FIG. 2. Identical parts are provided with the same reference symbols in all the figures.
- the gas turbine 1 has a compressor 2 for combustion air, a combustion chamber 4 and a turbine 6 for driving the compressor 2 and a generator (not shown) or a work machine.
- the turbine 6 and the compressor 2 are arranged on a common turbine shaft 8, also referred to as a turbine rotor, to which the generator or the working machine is also connected, and which is rotatably mounted about its central axis 9.
- the combustion chamber 4 which is designed as an annular combustion chamber, is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel. It is also provided on its inner wall with heat shield elements, not shown.
- the turbine 6 has a number of rotatable rotor blades 12 connected to the turbine shaft 8.
- the blades 12 are arranged in a ring shape on the turbine shaft 8 and thus form a number of rows of blades.
- the turbine 6 comprises a number of stationary guide vanes 14, which are also attached to an inner casing 16 of the turbine 6 in a ring shape, with the formation of rows of guide vanes.
- the blades 12 are used to drive the turbine shaft 8 by transfer of momentum from the working medium M flowing through the turbine 6.
- the guide blades 14, serve to guide the flow of the working medium M between two successive rows of blades or rotor blades as seen in the flow direction of the working medium M.
- a successive pair of a ring of guide vanes 14 or a row of guide vanes and a ring of rotor blades 12 or a row of rotor blades is also referred to as a turbine stage.
- Each guide vane 14 has a platform 18, also referred to as a blade root, which is arranged as a wall element for fixing the respective guide vane 14 to the inner housing 16 of the turbine 6.
- the platform 18 is a thermally comparatively heavily loaded component, which forms the outer boundary of a heating gas channel for the working medium M flowing through the turbine 6.
- Each rotor blade 12 is fastened in an analogous manner to the turbine shaft 8 via a platform 20 which is also referred to as a blade root.
- each guide ring 21 is arranged on the inner casing 16 of the turbine 6.
- the outer surface of each guide ring 21 is likewise exposed to the hot working medium M flowing through the turbine 6 and is spaced in the radial direction from the outer end 22 of the rotor blade 12 lying opposite it by a gap.
- the guide rings 21 arranged between adjacent guide vane rows serve in particular as cover elements which protect the inner wall 16 or other housing installation parts against thermal overloading by the hot working medium M flowing through the turbine 6.
- the gas turbine 1 is designed for a comparatively high outlet temperature of the working medium M emerging from the combustion chamber 4 of approximately 1200 ° C. to 1500 ° C.
- its essential components, in particular the combustion chamber 4 are designed to be coolable.
- the combustion chamber wall 23 is designed as a tube construction and is constructed from a large number of coolant tubes 24 which are connected to one another in a gas-tight manner to form the combustion chamber wall 23.
- the combustion chamber 4 is designed as a so-called annular combustion chamber, in which a plurality of burners 10 arranged in the circumferential direction around the turbine shaft 8 open into a common combustion chamber space.
- the combustion chamber 4 is in its entirety as an annular
- the combustion chamber 4 has an initial or inflow section into which the outlet of the respectively assigned burner 10 opens.
- the cross section of the combustion chamber 4 then narrows, the resulting flow profile of the working medium M being taken into account in this area.
- the combustion chamber 4 On the output side, the combustion chamber 4 has a curvature in longitudinal section, by means of which the outflow of the working medium M from the combustion chamber 4 is favored in a first rotor blade row, which is seen for a particularly high impulse and energy transfer to the downstream flow side.
- the combustion chamber wall 23 is formed both in the outer region of the combustion chamber 4 and in the inner region thereof by coolant tubes 24, the longitudinal axis of which is essentially parallel to the flow direction of the working medium M in the interior of the
- Combustion chamber 4 are aligned.
- the coolant tubes 24 are made of cast material, which is selected in particular in view of the particularly high mechanical and thermal strength of the coolant tubes.
- each coolant tube 24 is formed in the exemplary embodiment by a suitable combination of a plurality of successive tube segments 26.
- the type and number of pipe segments 26 are chosen such that, on the one hand, with regard to the length and shape of each pipe segment 26 and with regard to the casting material used, a particularly high mechanical strength of each individual pipe segment 26 is ensured, while on the other hand the shape also in each case un - Taking into account the desired flow path for the working medium M is selected appropriately.
- the comparatively strong local curvatures that may be desired can be provided in a particularly simple and reliable manner by the segmentation of the coolant tubes 24.
- the coolant tubes 24 are also designed for particular strength, particularly with regard to locally varying thermal loads and the resulting thermal stresses.
- the coolant tubes 24 and in particular the tube segments 26 forming them are essentially trapezoidal in cross-section, as is shown for the middle piece of a tube segment 26 in FIG. 3a.
- the coolant tubes 24 have a comparatively longer inside 28 and a comparatively shorter outside 30 in cross section.
- a suitable seal for example a brush seal seal 32, is provided for sealing the interspaces between adjacent coolant tubes 24, so that the suitable combination of the coolant tubes 24 results in a closed and sealed combustion chamber 4 on the gas side.
- the trapezoidal design of the pipe cross sections favors in particular a flat design of the structure obtainable by joining adjacent coolant pipes 24, so that the closed expansion management of the combustion chamber 4 can be reached in a comparatively simple manner.
- a connection which is particularly simple with regard to assembly or maintenance purposes, is provided between two tube segments 26 of each coolant tube 24 that follow one another on the coolant side.
- successive tube segments 26 of a coolant tube 24 are connected to one another via an associated transition piece 34.
- each tube segment 26 is essentially round in its cross-section in its end regions to form the respective transition piece 34, as shown in FIG. 3b.
- the production of the coolant tubes 24 from cast material enables the respective transition piece 34 to be molded onto the respective tube segment 26 in a comparatively simple manner, with the actually trapezoidal cross section of the respective tube segment 26 being continuously transferred to the circular cross section provided at the end in the transition region.
- the respective transition pieces 34 are shifted into the outer region of the combustion chamber 4 with regard to their central line and in comparison to the middle pieces of the respective pipe segments 26, so that under
- the coolant tubes 24 are fastened to a number of common carrier rings 36 which, viewed in the longitudinal direction or in the flow direction of the working medium M, enclose the combustion chamber 4 formed from the actual coolant tubes 24 in a suitably selected spacing.
- the respective coolant tubes 24 or the tube segments 26 forming these are on the carrier rings 36 via coolable screws 38 attached, as shown in the embodiment of Figure 3c.
- the carrier rings 36 are connected to one another by longitudinal ribs oriented essentially in the longitudinal direction or in the flow direction of the working medium M.
- the design of the combustion chamber 4 as a tubular construction makes it possible to apply a comparatively large amount of cooling air as coolant K to the combustion chamber wall 23 with only comparatively small pressure losses.
- the coolant K emerging from the coolant tubes 24 is to be fed into the as exclusive or additional combustion air Combustion chamber 4 is provided.
- a supply of the coolant K to the coolant pipes 24 is provided at their end assigned to the outlet of the combustion chamber 4.
- the coolant K is fed to the coolant tubes 24 there, as can be seen in FIG. 2, via suitable inflow openings 42.
- the inflow openings 42 are positioned with respect to their spatial alignment in such a way that in the outlet area of the combustion chamber 4 the impinging cooling of the respective pipe segment 26 takes place first due to the cooling air flowing in as coolant K. Subsequently, the coolant K is deflected within the respective pipe segment 26, and then the coolant K flows through the respective coolant pipe 24 in its longitudinal direction, the cooling taking place by contact of the coolant K with the respective pipe wall.
- the respective burner 10 from the outlet area of the combustion chamber 4 to its confluence area, in which the respective burner 10 is also arranged.
- the coolant K which is now heated or preheated by the continuous cooling of the respective coolant tube 24 flows out of the coolant tubes 24 and is then assigned to a respective downstream collecting space 46.
- the coolant tubes 24 are connected on the output side to the respectively assigned burner 10 via this collecting space 46, so that the coolant K flowing out of the coolant tubes 24 can be used as combustion air in the respective burner 10.
- the feeding of the respective burner 10 with combustion air can be provided exclusively via the coolant K flowing out of the coolant tubes 24 or also with additional combustion air which may be additionally required and is supplied externally.
- the combustion chamber 4 is designed as an annular combustion chamber, it is usually advantageous to arrange the burners 10 as symmetrically as possible and consequently to set the flow conditions within the combustion chamber 4 as symmetrically as possible.
- This principle is also taken into account in the gas turbine 1 on the coolant side, with in particular each burner 10 being assigned the same number of coolant tubes 24 on the combustion air side.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Priority Applications (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP03798881A EP1537363A1 (fr) | 2002-09-13 | 2003-09-01 | Turbine a gaz |
US10/525,780 US20050247062A1 (en) | 2002-09-13 | 2003-09-01 | Gas turbine |
JP2004540568A JP4181546B2 (ja) | 2002-09-13 | 2003-09-01 | ガスタービン |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP02020694A EP1398569A1 (fr) | 2002-09-13 | 2002-09-13 | Turbine à gaz |
EP02020694.2 | 2002-09-13 |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2004031656A1 true WO2004031656A1 (fr) | 2004-04-15 |
Family
ID=31725437
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/EP2003/009703 WO2004031656A1 (fr) | 2002-09-13 | 2003-09-01 | Turbine a gaz |
Country Status (5)
Country | Link |
---|---|
US (1) | US20050247062A1 (fr) |
EP (2) | EP1398569A1 (fr) |
JP (1) | JP4181546B2 (fr) |
CN (1) | CN100394110C (fr) |
WO (1) | WO2004031656A1 (fr) |
Families Citing this family (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8695989B2 (en) | 2004-04-30 | 2014-04-15 | Siemens Aktiengesellschaft | Hot gas seal |
GB2434199B (en) * | 2006-01-14 | 2011-01-05 | Alstom Technology Ltd | Combustor liner with heat shield |
EP1862740B1 (fr) * | 2006-05-31 | 2015-09-16 | Siemens Aktiengesellschaft | Paroi de chambre de combustion |
US8397512B2 (en) * | 2008-08-25 | 2013-03-19 | General Electric Company | Flow device for turbine engine and method of assembling same |
EP2405200A1 (fr) * | 2010-07-05 | 2012-01-11 | Siemens Aktiengesellschaft | Appareil de combustion et moteur de turbine à gaz |
US9534783B2 (en) * | 2011-07-21 | 2017-01-03 | United Technologies Corporation | Insert adjacent to a heat shield element for a gas turbine engine combustor |
DE102011083814A1 (de) * | 2011-09-30 | 2013-04-04 | Mtu Aero Engines Gmbh | Segmentiertes Bauteil |
WO2015017180A1 (fr) * | 2013-08-01 | 2015-02-05 | United Technologies Corporation | Système de fixation pour panneau de cloison en céramique |
CN104454174A (zh) * | 2014-10-13 | 2015-03-25 | 罗显平 | 一种提高燃气发动机动力输出功率的方法 |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB376974A (en) * | 1930-09-02 | 1932-07-21 | Bbc Brown Boveri & Cie | Improvements in and relating to combustion chambers |
FR980028A (fr) * | 1942-06-18 | 1951-05-07 | Regent | Perfectionnements apportés aux chambres de combustion |
DE1025915B (de) * | 1953-07-03 | 1958-03-13 | Still Fa Carl | Gasbeheizter Roehrenerhitzer mit einem aus Rohren gebildeten selbsttragenden Feuerraum |
US3398527A (en) * | 1966-05-31 | 1968-08-27 | Air Force Usa | Corrugated wall radiation cooled combustion chamber |
US5129447A (en) * | 1991-05-20 | 1992-07-14 | United Technologies Corporation | Cooled bolting arrangement |
DE4343332A1 (de) * | 1993-12-20 | 1995-06-22 | Abb Management Ag | Vorrichtung zur Konvektivkühlung einer dichten Brennkammer |
US5832718A (en) * | 1995-12-19 | 1998-11-10 | Daimler-Benz Aerospace Airbus Gmbh | Combustion chamber especially for a gas turbine engine using hydrogen as fuel |
Family Cites Families (30)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US1935659A (en) * | 1930-09-01 | 1933-11-21 | Bbc Brown Boveri & Cie | Pressureproof combustion chamber |
US3190070A (en) * | 1950-04-05 | 1965-06-22 | Thiokol Chemical Corp | Reaction motor construction |
US3043103A (en) * | 1958-10-10 | 1962-07-10 | Gen Motors Corp | Liquid cooled wall |
US3066702A (en) * | 1959-05-28 | 1962-12-04 | United Aircraft Corp | Cooled nozzle structure |
US3031844A (en) * | 1960-08-12 | 1962-05-01 | William A Tomolonius | Split combustion liner |
US3177935A (en) * | 1963-12-17 | 1965-04-13 | Irwin E Rosman | Cooling tube structure |
DE2015024B2 (de) * | 1970-03-28 | 1971-10-14 | Messerschmitt Bolkow Blohm GmbH, 8000 München | Verfahren zur herstellung von regenerativ gekuehlten brenn kammern und oder schubduesen |
US4288980A (en) * | 1979-06-20 | 1981-09-15 | Brown Boveri Turbomachinery, Inc. | Combustor for use with gas turbines |
US4765145A (en) * | 1987-01-20 | 1988-08-23 | Rockwell International Corporation | Connector assembly |
US5024058A (en) * | 1989-12-08 | 1991-06-18 | Sundstrand Corporation | Hot gas generator |
US5636508A (en) * | 1994-10-07 | 1997-06-10 | Solar Turbines Incorporated | Wedge edge ceramic combustor tile |
JPH08270950A (ja) * | 1995-02-01 | 1996-10-18 | Mitsubishi Heavy Ind Ltd | ガスタービン燃焼器 |
US5832719A (en) * | 1995-12-18 | 1998-11-10 | United Technologies Corporation | Rocket thrust chamber |
DE19751299C2 (de) * | 1997-11-19 | 1999-09-09 | Siemens Ag | Brennkammer sowie Verfahren zur Dampfkühlung einer Brennkammer |
DE19804232C2 (de) * | 1998-02-04 | 2000-06-29 | Daimler Chrysler Ag | Brennkammer für Hochleistungstriebwerke und Düsen |
JP4271866B2 (ja) * | 1998-11-27 | 2009-06-03 | ボルボ エアロ コーポレイション | 冷却されるノズル壁を有するロケットノズルのノズル構造 |
DE19915082C1 (de) * | 1999-04-01 | 2000-07-13 | Daimler Chrysler Ag | Verfahren zur Herstellung einer gekühlten Düse für ein Raketentriebwerk |
WO2002055862A1 (fr) * | 2001-01-11 | 2002-07-18 | Volvo Aero Corporation | Element de moteur-fusee et procede de fabrication d'un element de moteur-fusee |
US20020157400A1 (en) * | 2001-04-27 | 2002-10-31 | Siemens Aktiengesellschaft | Gas turbine with combined can-type and annular combustor and method of operating a gas turbine |
RU2289035C2 (ru) * | 2001-12-18 | 2006-12-10 | Вольво Аэро Корпорейшн | Подверженный во время работы воздействию высоких тепловых нагрузок элемент конструкции и способ его изготовления |
ES2285129T3 (es) * | 2002-05-28 | 2007-11-16 | Volvo Aero Corporation | Estructura de pared. |
EP1389714A1 (fr) * | 2002-08-16 | 2004-02-18 | Siemens Aktiengesellschaft | Chambre à combustion de turbine à gaz |
ES2307834T3 (es) * | 2003-01-29 | 2008-12-01 | Siemens Aktiengesellschaft | Camara de combustion. |
US6931855B2 (en) * | 2003-05-12 | 2005-08-23 | Siemens Westinghouse Power Corporation | Attachment system for coupling combustor liners to a carrier of a turbine combustor |
EP1482246A1 (fr) * | 2003-05-30 | 2004-12-01 | Siemens Aktiengesellschaft | Chambre de combustion |
US7213392B2 (en) * | 2003-06-10 | 2007-05-08 | United Technologies Corporation | Rocket engine combustion chamber |
US7146815B2 (en) * | 2003-07-31 | 2006-12-12 | United Technologies Corporation | Combustor |
US7043921B2 (en) * | 2003-08-26 | 2006-05-16 | Honeywell International, Inc. | Tube cooled combustor |
US7370469B2 (en) * | 2004-12-13 | 2008-05-13 | United Technologies Corporation | Rocket chamber heat exchanger |
US7721547B2 (en) * | 2005-06-27 | 2010-05-25 | Siemens Energy, Inc. | Combustion transition duct providing stage 1 tangential turning for turbine engines |
-
2002
- 2002-09-13 EP EP02020694A patent/EP1398569A1/fr not_active Withdrawn
-
2003
- 2003-09-01 EP EP03798881A patent/EP1537363A1/fr not_active Withdrawn
- 2003-09-01 US US10/525,780 patent/US20050247062A1/en not_active Abandoned
- 2003-09-01 CN CNB038215306A patent/CN100394110C/zh not_active Expired - Fee Related
- 2003-09-01 JP JP2004540568A patent/JP4181546B2/ja not_active Expired - Fee Related
- 2003-09-01 WO PCT/EP2003/009703 patent/WO2004031656A1/fr active Application Filing
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB376974A (en) * | 1930-09-02 | 1932-07-21 | Bbc Brown Boveri & Cie | Improvements in and relating to combustion chambers |
FR980028A (fr) * | 1942-06-18 | 1951-05-07 | Regent | Perfectionnements apportés aux chambres de combustion |
DE1025915B (de) * | 1953-07-03 | 1958-03-13 | Still Fa Carl | Gasbeheizter Roehrenerhitzer mit einem aus Rohren gebildeten selbsttragenden Feuerraum |
US3398527A (en) * | 1966-05-31 | 1968-08-27 | Air Force Usa | Corrugated wall radiation cooled combustion chamber |
US5129447A (en) * | 1991-05-20 | 1992-07-14 | United Technologies Corporation | Cooled bolting arrangement |
DE4343332A1 (de) * | 1993-12-20 | 1995-06-22 | Abb Management Ag | Vorrichtung zur Konvektivkühlung einer dichten Brennkammer |
US5832718A (en) * | 1995-12-19 | 1998-11-10 | Daimler-Benz Aerospace Airbus Gmbh | Combustion chamber especially for a gas turbine engine using hydrogen as fuel |
Also Published As
Publication number | Publication date |
---|---|
EP1537363A1 (fr) | 2005-06-08 |
EP1398569A1 (fr) | 2004-03-17 |
CN100394110C (zh) | 2008-06-11 |
JP2005538310A (ja) | 2005-12-15 |
JP4181546B2 (ja) | 2008-11-19 |
US20050247062A1 (en) | 2005-11-10 |
CN1682078A (zh) | 2005-10-12 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP1636526B1 (fr) | Chambre a combustion | |
EP2342427B1 (fr) | Support d'aubes statorique axialement segmenté d'une turbine à gaz | |
EP1451450B1 (fr) | Ensemble turbine a gaz | |
EP1443275A1 (fr) | Chambre de combustion | |
EP1724526A1 (fr) | Coquille de turbine à gaz, turbine à gaz et procédé de démarrage et d'arrêt d'une turbine à gaz | |
WO2004031656A1 (fr) | Turbine a gaz | |
DE102019104814B4 (de) | Mit einem Einsatzträger ausgestattete Turbinenschaufel | |
EP2347101B1 (fr) | Turbine à gaz et moteur à turbine à gaz associé | |
EP2206885A1 (fr) | Turbine à gaz | |
WO2009109430A1 (fr) | Dispositif d’étanchéité et turbine à gaz | |
EP1731715A1 (fr) | Transition d'une chambre de combustion à une turbine | |
EP2347100B1 (fr) | Turbine à gaz avec insert de refroidissement | |
EP2196628A1 (fr) | Support d'aube directrice | |
EP1429077B1 (fr) | Turbine à gaz | |
EP2218882A1 (fr) | Système de support d'aube directrice | |
DE19544011B4 (de) | Strömungsmaschine | |
EP1529181B1 (fr) | Chambre de combustion de turbine a gaz | |
EP1422479B1 (fr) | Chambre pour la combustion d' un mélange combustible fluide | |
WO2004109187A1 (fr) | Element de protection thermique | |
EP1564376B1 (fr) | Construction de rotor pour turbomachine | |
EP1588102A2 (fr) | Element de bouclier thermique, chambre de combustion et turbine a gaz | |
WO2010023150A1 (fr) | Support d'aubes directrices pour une turbine à gaz | |
EP2194236A1 (fr) | Carter de turbine | |
EP2184449A1 (fr) | Support d'aube directrice, turbine à gaz et moteur à turbine à gaz ou à vapeur avec un tel support d'aube directrice | |
EP1284392A1 (fr) | Chambre de combustion |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AK | Designated states |
Kind code of ref document: A1 Designated state(s): CN JP US |
|
AL | Designated countries for regional patents |
Kind code of ref document: A1 Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IT LU MC NL PT RO SE SI SK TR |
|
121 | Ep: the epo has been informed by wipo that ep was designated in this application | ||
WWE | Wipo information: entry into national phase |
Ref document number: 10525780 Country of ref document: US |
|
WWE | Wipo information: entry into national phase |
Ref document number: 2003798881 Country of ref document: EP |
|
WWE | Wipo information: entry into national phase |
Ref document number: 20038215306 Country of ref document: CN Ref document number: 2004540568 Country of ref document: JP |
|
WWP | Wipo information: published in national office |
Ref document number: 2003798881 Country of ref document: EP |